Hybrid turbine nozzle

ABSTRACT

A turbine nozzle apparatus for a gas turbine engine includes: an annular inner band; an annular outer band circumscribing the inner band; a plurality of airfoil-shaped structural vanes extending between and interconnecting the inner band and the outer band; and a plurality of airfoil-shaped non-structural vanes extending between the inner band and the outer band, each non-structural vane having a root end received by the inner band and a tip end received by the outer band, such that each non-structural vane is free to move to a limited degree relative to the inner and outer bands.

BACKGROUND OF THE INVENTION

This invention relates generally to gas turbine engines, and more particularly to turbine nozzles for such engines incorporating airfoils made of a low-ductility material.

A typical gas turbine engine includes a turbomachinery core having a high pressure compressor, a combustor, and a high pressure turbine in serial flow relationship. The core is operable in a known manner to generate a primary gas flow. The high pressure turbine (also referred to as a gas generator turbine) includes one or more stages which extract energy from the primary gas flow. Each stage comprises a stationary turbine nozzle followed by a downstream rotor carrying turbine blades. These components operate in an extremely high temperature environment, and must be cooled by air flow to ensure adequate service life. Typically, the air used for cooling is extracted (bled) from the compressor. Bleed air usage negatively impacts specific fuel consumption (“SFC”) and should generally be minimized.

Metallic turbine structures can be replaced with materials having better high-temperature capabilities, such as ceramic matrix composites (“CMCs”). The density of CMCs is approximately one-third of that of conventional metallic superalloys used in the hot section of turbine engines, so by replacing the metallic alloy with CMC while maintaining the same airfoil geometry, the weight of the component decreases. By replacing a majority of the airfoils in a turbine nozzle, the total weight of the assembly decreases, as well as the need for cooling air flow.

CMC and similar materials have unique mechanical properties that must be considered during design and application of an article such as a shroud segment. For example, CMC materials have relatively low tensile ductility or low strain to failure when compared with metallic materials. Also, CMCs have a coefficient of thermal expansion (“CTE”) approximately one-third that of superalloys, which means that a rigid joint between the two different materials induces large strains and stresses with a change in temperature from the assembled condition. The allowable stress limits for CMCs are also lower than metal alloys which drives a need for simple and low stress design for CMC components.

Accordingly, there is a need for an apparatus for mounting CMC and other low-ductility airfoils that minimizes mechanical loads on those components.

BRIEF DESCRIPTION OF THE INVENTION

This need is addressed by the present invention, which provides a turbine nozzle including non-structural airfoils which are positioned and retained to a surrounding structure while permitting limited freedom of movement.

According to one aspect of the invention, a shroud apparatus for a gas turbine engine includes a turbine nozzle apparatus for a gas turbine engine includes: an annular inner band; an annular outer band circumscribing the inner band; a plurality of airfoil-shaped structural vanes extending between and interconnecting the inner band and the outer band; and a plurality of airfoil-shaped non-structural vanes extending between the inner band and the outer band, each non-structural vane having a root end received by the inner band and a tip end received by the outer band, such that each non-structural vane is free to move to a limited degree relative to the inner and outer bands.

According to another aspect of the invention, a method is provided for assembling a turbine nozzle for a gas turbine engine. The method includes: providing an annular inner band, an annular outer band circumscribing the inner band, and a plurality of airfoil-shaped structural vanes extending between and interconnecting the inner band and the outer band; inserting an airfoil-shaped non-structural vane through an opening formed in one of the inner and outer bands; and closing the opening, such that the non-structural vanes extend between the inner band and the outer band, each non-structural vane having a root end received by the inner band and a tip end received by the outer band, such that each non-structural vane is free to move to a limited degree relative to the inner and outer bands.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which:

FIG. 1 is a schematic perspective view of a turbine nozzle assembly for a gas turbine engine, constructed according to an aspect of the present invention;

FIG. 2 is an enlarged view of a portion of the turbine nozzle shown in FIG. 1;

FIG. 3 is a schematic perspective view of portion of an alternative turbine nozzle for a gas turbine engine;

FIG. 4 is an enlarged view of a portion of the turbine nozzle shown in FIG. 3;

FIG. 5 is a cross-sectional view of a portion of a turbine nozzle; and

FIG. 6 is a view taken along lines 6-6 of FIG. 5.

DETAILED DESCRIPTION OF THE INVENTION

Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views, FIGS. 1 and 2 depict an exemplary turbine nozzle 10 constructed according to an aspect of the present invention. The turbine nozzle 10 is a stationary component forming part of a turbine section of a gas turbine engine. It will be understood that the turbine nozzle 10 would be mounted in a gas turbine engine upstream of a turbine rotor with a rotor disk carrying an array of airfoil-shaped turbine blades, the nozzle and the rotor defining one stage of the turbine. The primary function of the nozzle is to direct the combustion gas flow into the downstream turbine rotor stage.

A turbine is a known component of a gas turbine engine of a known type, and functions to extract energy from high-temperature, pressurized combustion gases from an upstream combustor (not shown) and to convert the energy to mechanical work, which is then used to drive a compressor, fan, shaft, or other mechanical load (not shown). The principles described herein are equally applicable to turbofan, turbojet and turboshaft engines, as well as turbine engines used for other vehicles or in stationary applications.

It is noted that, as used herein, the term “axial” or “longitudinal” refers to a direction parallel to an axis of rotation of a gas turbine engine, while “radial” refers to a direction perpendicular to the axial direction, and “tangential” or “circumferential” refers to a direction mutually perpendicular to the axial and tangential directions. (See arrows “A”, “R”, and “T” in FIG. 1). These directional terms are used merely for convenience in description and do not require a particular orientation of the structures described thereby.

The turbine nozzle 10 includes an annular inner band 12 and an annular outer band 14, which define the inner and outer boundaries, respectively, of a hot gas flowpath through the turbine nozzle 10.

An array of airfoil-shaped turbine vanes is disposed between the inner band 12 and the outer band 14. The array of vanes includes a group of structural vanes 16A (marked with an “x” in FIG. 1) alternating with a group of non-structural vanes 16B. The turbine nozzle 10 may be considered a “hybrid” structure in that the structural and non-structural vanes 16A and 16B are made from materials with different properties.

Each structural vane 16A has opposed concave and convex sides extending between a leading edge and a trailing edge, and extends between a root end 18 and a tip end 20. A sufficient number of structural vanes 16A are provided so as to maintain a concentric relationship between the inner band 12 and the outer band 14 during engine operation and to control the relative thermal growth between the inner band 12 and the outer band 14. As used herein, the term “structural” identifies vanes 16A which are configured and mounted so as to transfer thermal and/or mechanical loads between the inner band 12 and the outer band 14. The structural vanes 16A are functionally integral with the inner and outer bands 12 and 14, and may be part of a single cast or forged component, or may be welded, brazed, or mechanically fastened to the inner and outer bands 12 and 14. In the illustrated example, the inner and outer bands 12 and 14 are each continuous annual structures, but alternatively one or both of the bands may be made up of a plurality of segments. In the specific example illustrated, there are 12 structural vanes 16A equally spaced around the circumference of the turbine nozzle 10.

The structural vanes 16A are constructed from a strong, ductile material such as a metal alloy. For example, a known type of nickel-, iron-, or cobalt-based “superalloy” may be used for this purpose.

Each non-structural vane 16B has opposed concave and convex sides extending between a leading edge and a trailing edge, and extends between a root end 22 and a tip end 24. One or more non-structural vanes 16B are disposed circumferentially between each pair of structural vanes 16A. In the specific example illustrated, there are 48 non-structural vanes 16B equally spaced around the circumference of the turbine nozzle 10, and the non-structural vanes 16B are disposed in groups of four. A single structural vane 16A separates adjacent groups of non-structural vanes 16B.

As used herein, the term “non-structural” identifies vanes 16B which are configured and mounted such that they do not transfer significant thermal and/or mechanical loads between the inner band 12 and the outer band 14.

It will be understood that all vanes 16A and 16B are individually subject to significant aerodynamic (e.g. gas pressure) loads, and must have sufficient stiffness and yield strength to withstand these loads in operation.

Each of the non-structural vanes 16B is constructed from a low-ductility, high-temperature-capable material One example of a suitable material for the non-structural vanes 16B is a ceramic matrix composite (CMC) material of a known type. Generally, commercially available CMC materials include a ceramic type fiber for example SiC, forms of which are coated with a compliant material such as Boron Nitride (BN). The fibers are carried in a ceramic type matrix, one form of which is Silicon Carbide (SiC). Typically, CMC type materials have a room temperature tensile ductility of no greater than about 1%, herein used to define and mean a low tensile ductility material. Generally CMC type materials have a room temperature tensile ductility in the range of about 0.4 to about 0.7%. This is compared with metals typically having a room temperature tensile ductility of at least about 5%, for example in the range of about 5 to about 15%.

In this example, the inner band 12 incorporates an array of airfoil-shaped blind root pockets 26 formed therein. Each root pocket 26 receives the root end 22 of one of the non-structural vanes 16B. The root pockets 26 are sized and shaped so that each permits a small gap between the root pocket 26 and the associated non-structural vane 16B.

The outer band 14 incorporates an array of apertures 28 formed therein. Each aperture 28 is centered between adjacent structural vanes 16A, and each structural vane 16A carries an outer band segment 30 at its tip end 20. An arcuate cover 32 is provided for each aperture 28. The covers 32 are sized and shaped such that when installed in the apertures 28, they form a continuous annular structure in cooperation with the outer band segments 30. In the illustrated example, each cover 32 has array of airfoil-shaped blind tip pockets 34 formed therein; alternatively, each cover 32 could have a fewer number of tip pockets 34, or an individual cover could be provided for each non-structural vane 16B. Each tip pocket 34 receives the tip end 24 of one of the non-structural vanes 16B. The tip pockets 34 are sized and shaped such that each permits a small gap between the tip pocket 34 and the associated non-structural vane 16B.

The turbine nozzle 10 is assembled as follows. First, the non-structural vanes 16B are inserted from radially outside the outer band 14, through the apertures 28, until their root ends 22 engage the root pockets 26. Next, a cover 32 is installed into each aperture 28. The tip ends 24 of the non-structural vanes 16B are then manipulated to enter the tip pockets 34 of the covers 32.

Finally, the covers 32 are secured in the apertures 28. This could be done, for example, using known brazing or welding techniques, or by using mechanical fasteners (not shown). After engine service, the covers 32 may optionally be removed, permitting the non-structural vanes 16B to be replaced as needed, without replacing the entire nozzle 10.

After assembly, the non-structural vanes 16B are “loosely” retained between the inner band 12 and the outer band 14 such that they are free to move to a predetermined, limited degree, for example about 0.25 mm (0.010 in.) to about 0 5 mm ( (0.020 in.). During engine operation, gas pressure on the non-structural vanes 16B loads them against the pockets 26 and 34, preventing further movement in longitudinal (axial) and tangential directions, while permitting the inner and outer bands 12 and 14 to move radially relative to the non-structural vanes 16B. It is noted that pins, tabs, holes or other similar features can be added in any combination required to more precisely control radial and axial location of the non-structural vanes 16B while still allowing free thermal growth between the structural and non-structural vanes. For example, FIGS. 5 and 6 illustrate a configuration in which a pin 33 passes transversely through the root end 22 of the nonstructural vane 16B and the root pocket 26, and a rib 35 formed as part of the tip pocket 34 engages a transverse slot 37 of the tip end 24 of the nonstructural vane 16B. The effect is to make the nonstructural vane 16B “ride” radially with the inner band 12 while allowing free radial motion relative to the outer band 14, and simultaneously preventing axial motion.

FIGS. 3 and 4 illustrate a portion of a turbine nozzle 110 similar in construction to the turbine nozzle 10 described above, illustrating variations in how the vanes may be mounted. The nozzle 110 includes an inner band 112, an outer band 114, and structural vanes 116A alternating with non-structural vanes 116B.

Each structural vane 116A has opposed concave and convex sides extending between a leading edge and a trailing edge, and extends between a root end 118 and a tip end 120. The number and position of structural vanes 116A is selected as described above. The structural vanes 116A are functionally integral with the inner and outer bands 112 and 114.

The structural vanes 116A are constructed from a strong, ductile material such as a metal alloy. For example, a known type of nickel-, iron-, or cobalt-based “superalloy” may be used for this purpose.

Each non-structural vane 116B has opposed concave and convex sides extending between a leading edge and a trailing edge, and extends between a root end 122 and a tip end 124. One or more non-structural vanes 116B are disposed circumferentially between each pair of structural vanes 116A.

Each of the non-structural vanes 116B is constructed from a low-ductility, high-temperature-capable material. One example of a suitable material for the non-structural vanes 116B is a CMC material as described above.

In this example, the outer band 114 incorporates an array of blind tip pockets 126 formed therein. Each tip pocket 126 receives the tip end 124 of one of the non-structural vanes 116B. The tip pockets 126 are sized and shaped so that each permits a small gap between the tip pocket 126 and the associated non-structural vane 116B.

The inner band 112 incorporates an array of split sections formed therein. Each split section is positioned between adjacent structural vanes 116A, and includes a fixed forward segment 128 that is functionally and structurally integral to the adjoining portions of the inner band 112, and a discrete aft segment 130. The forward segment 128 and the aft segment 130 meet along a splitline “S” that lies in a radial-tangential plane. The forward segment 128 defines forward sections 132 of a plurality of airfoil-shaped root pockets 134, and the aft segment 130 defines aft sections 136 of the root pockets 134.

The aft segments 130 are sized and shaped such that when installed against the forward segments 128, they form a continuous annular structure in cooperation with the forward segments 128. This configuration of forward and aft segments may be referred herein to as a “split band” configuration. Each root pocket 134 receives the root end 122 of one of the non-structural vanes 116B. The root pockets 134 are sized and shaped such that each permits a small gap between itself and the associated non-structural vane 116B. The forward section 132 of the root pocket 134 may be made deeper in a radial direction than the aft section 136, to facilitate installation of the non-structural vanes 116B.

The turbine nozzle 110 is assembled as follows. First, the non-structural vanes 116B are inserted from radially inside the outer band 114, until their tip ends 124 engage the tip pockets 126. The root ends 122 of the non-structural vanes 116B are pivoted into the forward sections 132 of the root pockets 134. The aft segments 130 are then installed with the aft sections 136 of the root pockets 134 receiving the root ends 122 of the nonstructural vanes 116B.

Finally, the aft segments 130 are secured to the forward segments 128. This could be done, for example, using known brazing or welding techniques, or by using mechanical fasteners (not shown). After engine service, the aft segments 130 may optionally be removed, permitting the non-structural vanes 116B to be replaced as needed, without replacing the entire nozzle 110.

It is noted that the configuration of the inner and outer bands may be varied as required to suit a particular application, so long as one of the two bands includes one of the features described above permitting assembly of the non-structural vanes into the turbine nozzle. In other words, one of the two bands of a turbine nozzle would include apertures and associated covers, or a fore/aft split structure. The other of the two bands could include only blind pockets, apertures and associated covers, or a fore/aft split structure.

The turbine nozzle described above has several advantages compared to the prior art. The turbine nozzle of the present invention has a lower weight as compared to a completely-metallic turbine nozzle, by using a majority of CMC airfoils within a metallic frame. This turbine nozzle can also work to reduce cooling flow, because the majority of airfoils do not require air cooling.

This configuration allows the metal frame to dictate the thermal growth response of the nozzle, while the CMC airfoils are free thermally to grow and carry only aerodynamic pressure loading. The CMC airfoils are seated to the inner and outer bands under running conditions by the aerodynamic loading, and the metallic bands and airfoil struts transfer the load to the outer case to allow conventional cantilevered nozzle configuration. The invention maintains very similar thermal response of the nozzle assembly to the rest of the engine, compared to a completely-metallic nozzle. Other features of a cantilevered nozzle (e.g. seals and shields) can be attached to this composite assembly in the same fashion as a full metallic nozzle.

The foregoing has described a turbine nozzle for a gas turbine engine. While specific embodiments of the present invention have been described, it will be apparent to those skilled in the art that various modifications thereto can be made without departing from the spirit and scope of the invention. Accordingly, the foregoing description of the preferred embodiment of the invention and the best mode for practicing the invention are provided for the purpose of illustration only and not for the purpose of limitation. 

What is claimed is:
 1. A turbine nozzle apparatus for a gas turbine engine, comprising: an annular inner band; an annular outer band circumscribing the inner band; a plurality of airfoil-shaped structural vanes extending between and interconnecting the inner band and the outer band; and a plurality of airfoil-shaped non-structural vanes extending between the inner band and the outer band, each non-structural vane having a root end received by the inner band and a tip end received by the outer band, such that each non-structural vane is free to move to a limited degree relative to the inner and outer bands.
 2. The apparatus of claim 1 wherein one of the inner and outer bands includes a plurality of blind pockets, each blind pocket receiving one of the ends of the non-structural vanes.
 3. The apparatus of claim wherein one of the inner and outer bands includes: a plurality of spaced-apart apertures formed therein; and a cover received in each aperture, the aperture including at least one blind pocket, each blind pocket receiving one of the ends of the non-structural vanes.
 4. The apparatus of claim 1 wherein one of the inner and outer bands includes at least one split section comprising: a forward segment that is integral to adjacent sections of the band, the forward segment defining a forward section of at least one airfoil-shaped blind pocket; and an aft segment mating with the forward segment along a splitline, the aft segment defining an aft section at least one airfoil-shaped blind pocket, wherein the blind pocket receives one of the ends of the non-structural vanes.
 5. The apparatus of claim 1 wherein the non-structural vanes comprising a ceramic matrix composite.
 6. The apparatus of claim 1 wherein the structural vanes comprise a metallic alloy.
 7. The apparatus of claim 1 wherein: the inner band includes a plurality of blind root pockets, each root pocket receiving one of the root ends of one of the non-structural vanes; and the outer band includes: a plurality of spaced-apart apertures formed therein; and a cover received in each aperture, the aperture including a plurality of blind pockets, each blind pocket receiving one of the tip ends of one of the non-structural vanes.
 8. The apparatus of claim 1 wherein: the outer band includes a plurality of blind tip pockets, each tip pocket receiving one of the tip ends of one of the non-structural vanes; and the inner band includes at least one split section comprising: a forward segment that is integral to adjacent sections of the inner band, the forward segment defining forward sections of a plurality of airfoil-shaped blind pockets; and an aft segment mating with the forward segment along a splitline, the aft segment defining aft sections of a plurality of airfoil-shaped blind pockets, wherein each of the blind pockets receives one of the root ends of the non-structural vanes.
 9. The apparatus of claim 1 wherein the inner and outer bands are divided into a plurality of segments.
 10. A method of assembling a turbine nozzle for a gas turbine engine, comprising: providing an annular inner band, an annular outer band circumscribing the inner band, and a plurality of airfoil-shaped structural vanes extending between and interconnecting the inner band and the outer band; inserting an airfoil-shaped non-structural vane through an opening formed in one of the inner and outer bands; and closing the opening, such that the non-structural vanes extend between the inner band and the outer band, each non-structural vane having a root end received by the inner band and a tip end received by the outer band, such that each non-structural vane is free to move to a limited degree relative to the inner and outer bands.
 11. The method of claim 10 wherein one of the inner and outer bands includes a plurality of blind pockets, each blind pocket receiving one of the ends of the non-structural vanes.
 12. The method of claim 10 wherein one of the inner and outer bands includes: a plurality of spaced-apart apertures formed therein; and a cover received in each aperture, the aperture including at least one blind pocket, each blind pocket receiving one of the ends of the non-structural vanes.
 13. The method of claim 10 wherein one of the inner and outer bands includes at least one split section comprising: a forward segment that is integral to adjacent sections of the band, the forward segment defining a forward section of at least one airfoil-shaped blind pocket; and an aft segment mating with the forward segment along a splitline, the aft segment defining an aft section at least one airfoil-shaped blind pocket, wherein the blind pocket receives one of the ends of the non-structural vanes.
 14. The method of claim 10 wherein the non-structural vanes comprising a ceramic matrix composite.
 15. The method of claim 10 wherein the structural vanes comprise a metallic alloy.
 16. The method of claim 10 wherein: the inner band includes a plurality of blind root pockets, each root pocket receiving one of the root ends of one of the non-structural vanes; and the outer band includes: a plurality of spaced-apart apertures formed therein; and a cover received in each aperture, the aperture including a plurality of blind pockets, each blind pocket receiving one of the tip ends of one of the non-structural vanes.
 17. The method of claim 10 wherein: the outer band includes a plurality of blind tip pockets, each tip pocket receiving one of the tip ends of one of the non-structural vanes; and the inner band includes at least one split section comprising: a forward segment that is integral to adjacent sections of the inner band, the forward segment defining forward sections of a plurality of airfoil-shaped blind pockets; and an aft segment mating with the forward segment along a splitline, the aft segment defining aft sections of a plurality of airfoil-shaped blind pockets, wherein each of the blind pockets receives one of the root ends of the non-structural vanes.
 18. The method of claim 10 wherein the inner and outer bands are divided into a plurality of segments. 